(1) Field of the Invention
The present invention relates to a rotorcraft having a stabilizer device. The rotorcraft may in particular be a helicopter.
The invention thus lies in the narrow technical field of rotorcraft stabilizer devices, which devices are subjected to phenomena that do not appear on airplanes.
(2) Description of Related Art
Conventionally, an aircraft comprises an airframe extending longitudinally from a front end to a rear end and on either side of an anteroposterior plane of symmetry.
Furthermore, an aircraft sometimes includes stabilizer devices at its rear end for stabilizing certain movements of the aircraft.
The stabilizer devices include a stabilizer surface for stabilizing yaw movement of the aircraft.
Such a yaw movement stabilizer surface is generally referred to as a “tail fin”.
Furthermore, such stabilizer devices include stabilizer means for stabilizing pitching movements of the aircraft. Means for stabilizing pitching movements conventionally comprise at least one stabilizer surface presenting an angle of absolute value lying in the range 0° and plus or minus 90° relative to the anteroposterior plane of symmetry of the aircraft.
Such means for stabilizing pitching movements are sometimes referred to as a “horizontal tail plane”, or more simply “tail plane” below. The term “tail plane” is used more particularly when the stabilizer means are not necessarily horizontal. The term “pitching stabilizer means” is also used.
The pitching stabilizer means may comprise at least one airfoil surface passing right through the rear end of the aircraft in a transverse direction, or indeed it may include a non-through airfoil surface extending transversely from said rear end.
Such stabilizer devices perform an essential stabilizing role during forward flight of an airplane, but they can be penalizing for a rotorcraft.
An airplane usually has at least one wing carrying the airframe, and also a tail plane and a tail fin at the rear end of the airframe.
In contrast, the rotorcraft has at least one rotor providing lift and possibly also propulsion, which rotor is known as the “main rotor” for convenience. The airframe of a rotorcraft then extends for example in a vertical direction from a bottom portion having landing gear to a top portion carrying such a main rotor.
A helicopter type rotorcraft is thus provided with at least one main rotor providing the aircraft with at least some of its lift and propulsion.
Furthermore, a helicopter having a single main rotor is sometimes fitted with a tail rotor carried by a tail at its rear end. The tail rotor serves in particular to oppose the yaw torque exerted on the fuselage by the main rotor. Furthermore, the tail rotor serves to control movements in yaw of the helicopter.
The tail rotor of a helicopter is then either arranged within a tail fin in the context of a ducted tail rotor of the Fenestron® type, or else it is carried by the tail fin in the context of an unducted tail rotor.
As a result, a rotorcraft may have at least one main rotor and a tail rotor that interacts in harmful manner on stabilizer devices.
A rotary wing aircraft, and in particular a helicopter, can also perform hovering flight or flight at very low speed, namely at speeds of less than 70 knots (kt) for example.
During these stages of hovering or low-speed flight, such stabilizer devices can be found to be harmful.
Under such circumstances, when the tail fin carries the tail rotor, the air stream generated by the tail rotor can impact against the tail fin during the stages of hovering or low speed flight. The tail fin then blocks that air stream in part, thereby reducing the yaw moment exerted by the tail rotor on the airframe of the aircraft.
Under such circumstances, the power necessary for operating the tail rotor needs to be increased in order to compensate for the loss of efficiency caused by the tail fin.
That phenomenon which is sometimes referred to as the “tail fin blocking phenomenon”, does not happen on an airplane since an airplane does not have a tail rotor.
In order to limit this need for increased power, the trailing edge of the tail fin can be truncated. Nevertheless, the tail fin then becomes less effective in forward flight because of the reduction in its wing area.
Likewise, pitching stabilizer means are effective during a stage of cruising flight, with effectiveness increasing with increasing forward speed of the helicopter. Furthermore, the effectiveness of pitching stabilizer means is maximized by maximizing its wing area.
Nevertheless, the flow of air passing through the main rotor of a conventional helicopter in flight is deflected downwards, and in certain flight situations it comes to impact against the pitching stabilizer means, in particular when moving in translation at low speed or during hovering flight. This flow of air then exerts forces on the pitching stabilizer means that the pilot needs to compensate by operating flight controls.
Nevertheless, when flying conditions vary, the deflection of the air flow is also modified. Consequently, the forces exerted on the pitching stabilizer means by the air flow are also modified.
This phenomenon is referred to by the person skilled in the art as the “attitude hump”, and it does not occur on an airplane.
During a stage of transition between hovering flight and cruising flight, e.g. in the range 40 kt to 70 kt, the forces exerted by the air flow passing through the main rotor tend mainly to cause the tail plane to lose lift and impart a nose-up attitude to the helicopter by striking the pitching stabilizer means. This stage of flight is commonly referred to as the “transition stage” insofar as it is generally situated at low speed between a stage of hovering flight and a stage of cruising flight.
In order to balance the helicopter, the pilot must then use the stick for controlling the cyclic pitch of the blades of the main rotor in order to reduce the nose-up attitude of the helicopter.
The loss of lift generated by the pitching stabilizer means under such conditions is harmful for the performance of the aircraft. Furthermore, the nose-up movement of the aircraft is harmful for a pilot's visibility, particularly during a stage of approaching a landing area.
Furthermore, optimizing the pitching stabilizer means by maximizing its wing area accentuates the attitude hump.
Under such circumstances, using pitch stabilizer means of large wing area on a helicopter appears to be impossible without leading to an increase in the attitude hump phenomenon.
In this context, a rotorcraft is subjected to difficulties that are unknown to airplane manufacturers.
The design of stabilizer devices for a rotorcraft, and in particular for a helicopter, thus involves a compromise between the stage of flight when moving fast in translation and stages of flight when hovering or at low speeds.
To remedy the problem, stabilizer devices have a stationary airfoil surface and an airfoil surface that is movable in rotation relative to the stationary airfoil surface. The position of the movable surface relative to the stabilizer surface can then be controlled using at least one actuator.
The function of the movable airfoil surface is to modify the camber of the stabilizer device in order to modify its lift.
The actuator can be controlled by means of flight controls and/or by a computer.
Although advantageous, the main difficulty with that solution lies in the critical nature of the function and of the control forces to which the actuator is subjected.
A tail plane flap may be turned through an angle of 70° in hovering flight relative to a forward flight position. This angle may be incompatible with the operating range of an electric actuator because of the high levels of force to be delivered.
Under such circumstances, the device may comprise an electric actuator backed up by a hydraulic actuator, thereby making its architecture more complex.
Furthermore, a tail plane with an active rotary flap requires an actuator with a large bandwidth that is servo-controlled by a calculation closed loop. The problem posed by that type of architecture lies in finding an actuator that operates at high frequency.
Document FR 2 689 854 describes a helicopter tail fin. The tail fin has an airfoil surface. The tail fin then has a flap that is movable in rotation by being hinged to the trailing edge of the airfoil surface. The angle through which the flap is turned relative to a neutral position is a function of the collective pitch angle of the blades of a rotor of the aircraft and a function of the forward speed of the aircraft.
Furthermore, documents are known relating to a technical field that is remote from that of the invention, namely the technical field of airplanes. These documents are mentioned solely by way of illustration.
Document EP 2 371 707 B1 relates, according to its paragraph 13, to reducing the area of the tail fin of an airplane without reducing the ability of a flap of the tail fin to move in rotation in the presence of a large yaw moment, i.e. in the event of an engine failure, unbalance resulting from transporting external loads, gusts of wind, or flooding of a runway.
To this end, Document EP 2 371 707 A2 describes a tail fin having an airfoil surface. The tail fin then has a telescopic flap that is movable in rotation by being hinged to the trailing edge of the airfoil surface.
The tail fin blocking and attitude hump phenomena are thus not mentioned in that document.
Likewise, Document FR 2 911 113 describes an airplane tail plane.
That tail plane has a rotary flap hinged to a slider that moves in translation relative to a stationary surface. The flap slides in particular relative to the stationary surface in order to maximize the area of the tail fin during takeoff and landing, i.e. at low speed, and in order to minimize the area of the tail plane in cruising flight, i.e. at high speed.
That Document FR 2 911 113 shows a horizontal tail plane in a deployed position during stages of takeoff and landing and in a retracted position during cruising flight.
Those effects appear to be harmful to the tail fin blocking and attitude hump phenomena that are encountered in a rotorcraft.
Document US 2013/313355 describes pitching stabilizer means having at least one slot passing through the thickness of the pitching stabilizer means. At least one deflector separates two compartments in said slot.
Documents EP 2 409 917, EP 2 708 466, and EP 2 105 378 are also known.